Gas turbine

ABSTRACT

A gas turbine engine, in particular an aircraft engine, includes: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with a rear carrier bearing device radially between the planet carrier and a static structure on the input side of the gearbox device or a front carrier bearing device radially between the planet carrier and a static structure on the output side of the gearbox device.

The invention relates to a gas turbine with the features of claim 1.

Gas turbine engines, in particular aircraft engines with geared turbofanengines require a suitable support for shaft arrangements driving thegearbox and/or the propulsive fan. One shafting arrangement of a gearedturbofan engine is described in EP 3 144 486 A1.

This issue is addressed by a gas turbine with the features of claim 1.

The gas turbine comprises a turbine connected via an input shaft deviceto a gearbox device having a sun gear, a planet carrier having aplurality of planet gears attached thereto, and a ring gear. Typically,the gearbox device is driven by a low pressure or intermediate pressureturbine of the gas turbine, i.e. the sun gear is connected to the inputshaft device.

The gearbox device reduces the rotational speed from the turbine to thepropulsive fan towards the front of the gas turbine engine making theoverall engine more efficient. As will be described further below, thegearbox devices can have different designs.

Depending on the design of the gearbox device, the planet carrier or thering gear is connected to the propulsive fan via an output shaft deviceof the gearbox device. The output shaft device can comprise severalparts and is generally a hollow shaft with a cross-sectional shapeadapted to the load case and the available space within the engine.

There are two alternatives for a bearing device for the planet carrier.The first alternative comprises a rear carrier bearing device radiallybetween the planet carrier and a static structure on the input side ofthe gearbox device. The second alternative comprises a front carrierbearing device radially between the planet carrier and a staticstructure on the output side of the gearbox device. Those bearingdevices provide radial support for the propulsive fan and the planetcarrier.

The bearing devices may comprise more than one bearing. As will bedescribed below, the bearing devices can be positioned axially veryclose to the gearbox device.

The rear carrier bearing device or the front carrier bearing device canin one embodiment comprise at least one roller bearing. It is e.g. alsopossible to use a double roller bearing with two parallel rows.Furthermore, it is possible that the bearing device comprise bearingswhich are set apart a certain distance. Those bearings can be identical(e.g. all roller bearing) or they can have a different design.

As mentioned above, the front carrier bearing device can be axiallyadjacent to the gearbox device on the output side, in particular with anaxial distance measured from the centreline of the gearbox devicebetween 0, 1 and 2 times the inner radius of a seat element for thefront carrier bearing device and/or the rear carrier bearing device isaxially adjacent to the gearbox device on the input side, in particularwith an axial distance measured from the centreline of the gearboxdevice between 0.5 and 2 times the inner radius of a seat element forthe rear carrier bearing device. This means that e.g. the part of thebearing devices closest to the centreline of the gearbox device can bepositioned on the input side or the output side of the gearbox device.

Towards the front of the engine a fan shaft bearing system may beradially located between a fan shaft as part of the output shaft deviceand a static front cone structure, in particular the fan shaft bearingsystem may be axially positioned within the width of the propulsive fan.The static front cone structure—as an example for general staticstructure within the gas turbine—is relative at rest to the output shaftdevice. The loads of the fan shaft bearing system can be transmitted tothe static part. In one embodiment the fan shaft bearing system has anouter diameter between 0.05 to 0.20 the diameter of the propulsive fan,in particular between 0.1 and 0.15 times the diameter of the propulsivefan.

In a further embodiment, the planet carrier comprising the seat elementwhich is extending axially to the front and/or the rear of the gearboxdevice provides a radial seat for the front carrier bearing deviceand/or the rear carrier bearing device.

In one embodiment the carrier bearing system is axially adjacent to thegearbox device on the input or output side, in particular with an axialdistance measured from the centreline of the gearbox device between 0.1and 4 times the inner radius of the inter-shaft bearing system.

In a further embodiment, the planet carrier of the gearbox devicecomprises a seat element extending axially to the front and/or the rearof the gearbox device providing a radial seat for the inter-shaftbearing system and/or the carrier bearing system. The seat element canprovide the outer radial seat for the inter-shaft bearing system and theradial inner seat for the carrier bearing system. The seat element canbe connected to the planet carrier or in one piece with the planetcarrier.

Further to the rear of the engine an input shaft bearing system may beradially located between the input shaft device and a static rearstructure, the input shaft bearing system in particular may comprise atleast one roller bearing. As in the bearing devices or system describedabove, the input shaft bearing system can comprise more than one row ofbearings, the rows being identical or different. The rows can be axiallydistanced. Alternatively a ball bearing could be used at location of theinput shaft bearing system and a roller bearing in the inter-shaftbearing system.

The shape of the output shaft device can be adapted to spatialrequirements. For providing sufficiently mechanical properties,embodiments of the output shaft device can comprise at least one axialcross-section with a conical, sigmoidal or logarithmical shape. In onealternative the fan shaft can be directly attached to the carrier.

In a further embodiment the output shaft device comprises a curvic orspline coupling. The coupling could e.g. the form of a bellow shaft toachieve a decoupling of the bending between the output shaft and thegearbox device.

In one embodiment of the gas turbine the load path for force and/ortorque from the driving turbine to the propulsive fan extends from thedriving turbine to the propulsive fan via the input shaft device, thegearbox device and the output shaft device and a through shaft.

The through shaft may be coupled to the fan shaft at a forward end, andto the input shaft via an inter-shaft bearing at a rearward end, and maybe arranged to rotate with the fan shaft. Consequently, rearward axialloads from the fan can be carried by the through shaft, which can bereacted by forward axial loads by the turbine. Consequently, the overallloads experienced by shaft bearings is reduced. Furthermore, gearboxcomponents may not have to carry significant axial loads, therebyreducing stress on gearbox components, and allowing for a weight andvolume reduction.

The through shaft may have a greater flexibility than the fan shaft inat least shear and bending. Advantageously, misalignment between the fanand input shaft can be accommodated by the through shaft, withoutimposing excessive loads on the bearings.

The through shaft may be arranged to extend through an internal annulusof the sun gear. Advantageously, the system has good packing density,and a small volume.

In one embodiment the ring gear is rigidly connected to the static frontcone structure, as it is the case in epicyclic gearbox devices.

Furthermore, it is possible that the connection between the input shaftthe planet gears comprises a friction locking spline connection.

In another embodiment input shaft comprises flexibility means, inparticular at least one groove structure or convolutions. The inputshaft may have a greater flexibility than the fan shaft in at leastbending and shear.

Furthermore, in another embodiment the fan shaft is torsionally stiff,and may be stiff in bending and in shear.

It is also possible that e.g. the gearbox device comprises an epicyclicgearbox with the ring gear being fixed relative to the other parts ofthe gearbox device and the output shaft device being connected to theplanet carrier.

Alternatively, the gearbox device comprises a planetary gearbox in stararrangement with the planet carrier fixed relative to the other parts ofthe gearbox device and the output shaft device being connected to thering gear.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

The gas turbine engine comprises a gearbox device that receives an inputfrom the core shaft and outputs drive to the fan so as to drive the fanat a lower rotational speed than the core shaft. The input to thegearbox device may be directly from the core shaft, or indirectly fromthe core shaft, for example via a spur shaft and/or gear. The core shaftmay rigidly connect the turbine and the compressor, such that theturbine and compressor rotate at the same speed (with the fan rotatingat a lower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox device may be arranged to be driven by the core shaft thatis configured to rotate (for example in use) at the lowest rotationalspeed (for example the first core shaft in the example above). Forexample, the gearbox device may be arranged to be driven only by thecore shaft that is configured to rotate (for example in use) at thelowest rotational speed (for example only be the first core shaft, andnot the second core shaft, in the example above). Alternatively, thegearbox device may be arranged to be driven by any one or more shafts,for example the first and/or second shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18 or 18.5. The bypass ratio may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds). The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110 Nkg⁻¹ s, 105 N kg⁻¹ s, 100 N kg⁻¹ s, 95 N kg⁻¹ s, 90 N kg⁻¹ s, 85 N kg⁻¹s or 80 N kg⁻¹ s. The specific thrust may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400 K, 1450 K, 1500 K,1550 K, 1600 K or 1650 K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55° C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 shows schematically the upper half of a front section of anembodiment of a gas turbine with a drive train with an input shaftdevice of an epicyclic gearbox device, a gearbox device and an outputshaft device extending to a propulsive fan including a rear carrierbearing device;

FIG. 5 shows schematically the upper half of a front section of anembodiment of a gas turbine with a drive train with an input shaftdevice of an epicyclic gearbox device, a gearbox device and an outputshaft device extending to a propulsive fan including a front carrierbearing device.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, alow-pressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low-pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by thelow-pressure compressor 14 and directed into the high-pressurecompressor 15 where further compression takes place. The compressed airexhausted from the high-pressure compressor 15 is directed into thecombustion equipment 16 where it is mixed with fuel and the mixture iscombusted. The resultant hot combustion products then expand through,and thereby drive, the high-pressure and low-pressure turbines 17, 19before being exhausted through the nozzle 20 to provide some propulsivethrust. The high-pressure turbine 17 drives the high-pressure compressor15 by a suitable interconnecting shaft 27. The fan 23 generally providesthe majority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low-pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox device 30 is shown by way of example in greaterdetail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear38 comprise teeth about their periphery to intermesh with the othergears. However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearboxdevice 30 generally comprise at least three planet gears 32.

The epicyclic gearbox device 30 illustrated by way of example in FIGS. 2and 3 is of the planetary type, in that the planet carrier 34 is coupledto an output shaft via linkages 36, with the ring gear 38 fixed. Inanother embodiment the carrier and the output shaft can be manufacturedas one part. However, any other suitable type of epicyclic gearboxdevice 30 may be used. By way of further example, the epicyclic gearboxdevice 30 may be a star arrangement, in which the planet carrier 34 isheld fixed, with the ring (or annulus) gear 38 allowed to rotate. Insuch an arrangement the fan 23 is driven by the ring gear 38. By way offurther alternative example, the gearbox device 30 may be a differentialgearbox in which the ring gear 38 and the planet carrier 34 are bothallowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox device 30 in the engine10 and/or for connecting the gearbox device 30 to the engine 10. By wayof further example, the connections (such as the linkages 36, 40 in theFIG. 2 example) between the gearbox device 30 and other parts of theengine 10 (such as the input shaft 26, the output shaft and the fixedstructure 24) may have any desired degree of stiffness or flexibility.By way of further example, any suitable arrangement of the bearingsbetween rotating and stationary parts of the engine (for example betweenthe input and output shafts from the gearbox device 30 and the fixedstructures, such as the gearbox casing) may be used, and the disclosureis not limited to the exemplary arrangement of FIG. 2. For example,where the gearbox device 30 has a star arrangement (described above),the skilled person would readily understand that the arrangement ofoutput and support linkages and bearing locations would typically bedifferent to that shown by way of example in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox device may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 22meaning that the flow through the bypass duct 22 has its own nozzle thatis separate to and radially outside the core engine nozzle 20. However,this is not limiting, and any aspect of the present disclosure may alsoapply to engines in which the flow through the bypass duct 22 and theflow through the core 11 are mixed, or combined, before (or upstream of)a single nozzle, which may be referred to as a mixed flow nozzle. One orboth nozzles (whether mixed or split flow) may have a fixed or variablearea. Whilst the described example relates to a turbofan engine, thedisclosure may apply, for example, to any type of gas turbine engine,such as an open rotor (in which the fan stage is not surrounded by anacelle) or turboprop engine, for example.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

In FIG. 4 a schematic view of the front section of geared turbofanengine 10 is shown. The view axially extends from the propulsive fan 23in the front to the low-pressure compressor 14 towards the rear. Thelow-pressure compressor 14 is only shown symbolically to indicate therelative position of the drive train and its units.

The drive train comprises an input shaft device 50 (e.g. comprising theshaft 26 shown in FIG. 1), here driven by the not shown low-pressureturbine 19. The input shaft device 50 is connected to the sun gear 28 ofthe epicyclical gearbox device 30. The output of the gearbox device 30takes place via the planet carrier 34 which is connected with an outputshaft device 60 which has a portion acting as a fan shaft 61. Thatportion is rigidly connected with the propulsive fan 23. In analternative embodiment, the output shaft 60 can be replaced by a directconnection of the fan disk 61 to the carrier 34.

Therefore, the input torque is transmitted from the input shaft device50 to the sun gear 28 of the gearbox device 30, and to some extent tothe ring gear mount. The planet carrier 34 transmits the output torque(at a reduced rotational speed) to the output shaft 60 and eventually tothe propulsive fan 23.

The input shaft device 50 and the output shaft device 60 are here shownin a simplified manner. It is possible that the shape of the shaftdevices 50, 60 can be more complex and comprises more than one piece.

The shafting arrangement of the embodiment shown in FIG. 4 alsocomprises several bearing systems e.g. for taking the mechanical loadsor for locating the propulsive fan 23 and the gearbox device 30.

The first bearing to be described is rear carrier bearing device 90being positioned radially between the planet carrier 34 and a staticstructure 91. This rear carrier bearing device 90 here comprises oneroller bearing. In alternative embodiments, more than one roller bearing(e.g. double bearings, two bearings of different design) or otherbearing designs can be used. It is also possible that different bearingsof the rear carrier bearing device 90 are positioned at differentlocations.

The rear carrier bearing device 90 is, in this embodiment, axiallyadjacent to the gearbox device 30 on the input side. The axial distancebetween the rear carrier bearing device 70 to the gearbox device 30 cane.g. be between 0.5 and 2 times the inner radius of the inner radius ofa seat element 39 for the rear carrier bearing device 70. This could bein the range of 1 to 100 mm measured from the axial front side of therear carrier bearing device 70 to a centreline 41 of the gearbox device30.

The fan axial load is transferred via the fan-shaft bearing system 80(roller bearing), via the gearbox device 30 and into the input-shaftbearing 95 towards the rear. With this arrangement the supportstructures of the bearings can be reduced.

The radial inner seat of the rear carrier bearing 70 is on seat element39 extending axially to the rear of the gearbox device 30. A similarseat element 34 might be provided on the planet carrier 34 on the outputside of the gearbox 30.

On the output side of the gearbox device 30, the output shaft device 60only has one bearing system, a fan shaft bearing system 80. The radialinner seat of that bearing system is on the fan shaft 61, being a partof the output shaft device 60. The radial outer seat of the fan shaftbearing system 80 is connected to a static front cone structure 81. Inthe embodiment shown a roller bearing is used in the fan shaft bearingsystem 80. In alternative embodiments, more than one roller bearing(e.g. double bearings, two bearings of different design) or otherbearing designs can be used. It would be possible to install a ballbearing and transfer the axial load to the fan 13 via the static frontcone structure 81.

In the embodiment described herein the fan shaft bearing system 80 canhave an outer diameter between 0.05 to 0.2 times the diameter of thepropulsive fan 13. This range can be between 175 and 1250 mm.

In an alternative embodiment, the fan shaft bearing system 80 is notdirectly located underneath the propulsive fan 23. The fan shaft 61 thenextends to the front from the fan shaft bearing system 80. The fan shaft61 connects directly to the carrier 34.

The output shaft device 60 in the embodiment shown in FIG. 4 comprisesessentially a cylindrical section adjacent to the output side of thegearbox device 30 and under the propulsive fan 23 (i.e. the fan shaftsection 61). In between there is a conical section 62 linking the twocylindrical sections. Conical in this context means that the axialcross-section in this part of the output shaft device 60 is a straightline inclined radially inwards. In other embodiments this linkingsection can have different shapes than the conic shape in FIG. 4.

In the embodiment shown in FIG. 4 the static front cone structure 81 andthe static structure 91 form together one cavity around the gearboxdevice 30.

The ring gear 38 is rigidly connected to the static front cone structure81 but alternatively, it can be connected to a different static partwithin the engine 10.

In the embodiment shown in FIG. 4 there also a shaft extending from therear part of the engine 10 axially beyond the gearbox device 30 to thefront; the through shaft 63.

The load path for force and/or torque from the driving turbine 19, i.e.the low-pressure turbine 19 to the propulsive fan 23 extends via theinput shaft device 50, the through shaft 63, the gearbox device 30, theoutput shaft device 60.

In particular, the output shaft 60 comprises radially inner 63 and outer62 parts, with an annular space being provided in between. Both theinner 63 and outer 62 parts are coupled to the fan shaft 61, with onlythe outer part 62 being coupled to the gearbox output (i.e. the planetcarrier 34 in this embodiment).

The inner shaft comprises the through shaft 63. The through shaft 62 andshaft 62 are directly coupled to the fan shaft 61, such that each of theshafts 61, 62, 63 rotate together at the speed of the output of thegearbox 30. The through shaft extends through an inner annulus of thesun gear 28, toward the aft of the engine. The through shaft 63terminates in an inter-shaft bearing 64 in the form of a thrust bearing.The thrust bearing 64 is situated between the through shaft 63 and theinput shaft 26, and permits relatively rotation therebetween. Theinter-shaft bearing 64 is in the form of a thrust bearing, whichtransfers axial thrust load from the fan shaft 61 to the low pressureshaft 26. Consequently, axial load is transferred from the fan shaft 62to the low pressure shaft 26 via the through shaft 63.

This arrangement has several advantages. Firstly, torque and axialthrust loads are essentially provided through separate paths, withtorque loads being transmitted through the shaft 62 and gearbox 30,while thrust loads are transmitted through the through shaft 63.Consequently, thrust loads on the gearbox are significantly reduced.This results in lower loads on gearbox components, which simplifiesdesign, and may result in a lighter overall system. Secondly, theoverall thrust that must be reacted by the bearings is greatly reduced.

In a turbofan gas turbine engine, a large proportion of overall thrustis generated by the fan 23. This is particularly the case in a highbypass ratio engine. This thrust generates a rearward reaction force onthe fan 23, which must be carried by bearings. At the same time, a largeforward reaction force is generated by the low pressure turbine 19, dueto core flow therethrough. This problem becomes particularly difficultto solve as gas turbine engine size increases. In a conventional gasturbine engine, these loads are balanced to some extent by a directconnection between the fan and low pressure turbine by the low pressureshaft. However, in a geared engine, this connection is lost, withtorsional force instead being provided through the gearbox. A directconnection is not generally possible, in view of the differentrotational speeds of the fan shaft 61 and the low pressure turbine 26.In the present invention, direct connection between the low pressureturbine 26 and fan shaft 61 is restored by the provision of the throughshaft 63 and inter-shaft bearing 64, which allows thrust to betransferred to the low pressure shafts via the through shaft 63, therebyreducing the overall thrust that must be reacted by the bearings.

A further important consideration in geared gas turbine engine design isthe accommodation of loads applied to different components, and theresultant stress and misalignment on those components.

For instance, the inertia of the fan 23 may result in significantlateral and bending loads in flight. Where all of the components arerigid, this results in significant stress being imparted into thosecomponents, and components connected thereto. On the other hand, theseloads must be reacted, and torque must be transferred from the turbines19 to the fan 23.

Consequently, in the present arrangement, varying flexibility isprovided by the various components. In particular, the gearbox outputshaft 62 and fan shaft 61 are relatively stiff in the lateral andtorsional directions, i.e. in bending, shear and torsion. Consequently,movement of the fan 23 results in loads being transferred into thegearbox. On the other hand, both the sun input shaft 50 and throughshaft 63 are flexible, i.e. have a lower stiffness, relative to eitheror both of the fan shaft 61 and the output shaft 62, such that, in use,a lateral deflection of the fan 23 results in a larger bending movementof the sun input and through shafts 63, 64 than of the output shaft 62.Consequently, movement of the fan 23, gearbox 30 and shafts 62, 63, 64is isolated from the turbine 19. This is thought to be preferable toproviding flexibility in the output shaft 62 for several reasons.Firstly, in view of the large diameter and conical shape of the outputshaft, this component will naturally tend to have a high lateralstiffness. Consequently, taking measures to reduce this stiffness (suchas by providing undulations in the cross section of the shaft 62) mayresult in reduction torsional stiffness, which will result in reducedeffectiveness of torque transfer. Secondly, reduced bending and shearstiffness will necessarily result in reduced torsional stiffness. Sincethe through shaft carries substantially no torque, this reducedtorsional stiffness has fewer negative effects on the torque capacity ofthe system.

In FIG. 5 a variation of the embodiment shown in FIG. 4 is shown.Reference can be made to the respective description. Here no rearcarrier bearing device 90 is used but a front carrier bearing device 70which is connected to the static front cone.

The variant of FIG. 5 has several potential advantages over that in FIG.4. For instance, provision of a bearing 70 forward of the planet carrierensures that lateral loads from the fan 23 are transferred to staticstructure forward of the planet carrier 34, rather than being borne bythe planet carrier itself, in addition to the planet and sun gears.Consequently, stress on these components is reduced, which may result inincreased life, or reduced cost, weight or volume.

In the embodiments shown in FIGS. 4 and 5 the input shaft 50 is shownschematically as a straight shaft. It is possible that in an alternativeembodiment the input shaft comprises flexibility means such as groovesor meandering sections to provide a defined flexibility in the shaft.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine, comprising: a turbine connected via an inputshaft device to a gearbox device having a sun gear, a planet carrierhaving a plurality of planet gears attached thereto, and a ring gear,the sun gear is connected to the input shaft device, the planet carrieror the ring gear is connected to a propulsive fan via an output shaftdevice of the gearbox device, with a rear carrier bearing deviceradially between the planet carrier and a static structure on the inputside of the gearbox device or a front carrier bearing device radiallybetween the planet carrier and a static structure on the output side ofthe gearbox device.
 2. The gas turbine of claim 1, wherein the rearcarrier bearing device or the front carrier bearing device comprise atleast one roller bearing.
 3. The gas turbine of claim 1, wherein thefront carrier bearing device axially adjacent to the gearbox device onthe output side.
 4. The gas turbine of claim 1, wherein a fan shaftbearing system is radially located between a fan shaft as part of theoutput shaft device and a static structure.
 5. The gas turbine of claim4, wherein the fan shaft bearing system has an outer diameter between0.05 to 0.2 the diameter of the propulsive fan, in particular between0.1 and 0.15 times the diameter of the propulsive fan.
 6. The gasturbine of claim 1, the planet carrier comprises the seat elementextending axially to the front and/or the rear of the gearbox deviceproviding a radial seat for the front carrier bearing device and/or therear carrier bearing device.
 7. The gas turbine of claim 1, wherein aninput shaft bearing system is radially located between the input shaftdevice and a static structure.
 8. The gas turbine of claim 1, whereinthe output shaft device comprises at least one axial cross-section witha conical, sigmoidal or logarithmical shape.
 9. The gas turbine of claim1, wherein the output shaft device comprises a curvic or a splinecoupling.
 10. The gas turbine of claim 1, wherein the load path forforce and/or torque from the driving turbine to the propulsive fanextends via the input shaft device, the gearbox device and the outputshaft device and a through shaft.
 11. The gas turbine engine of claim10, wherein the through shaft is coupled to the fan shaft at a forwardend, and to the input shaft via an inter-shaft bearing at a rearwardend.
 12. The gas turbine engine of claim 10, wherein the through shafthas a greater flexibility than the fan shaft in at least shear andbending.
 13. The gas turbine engine of claim 10, wherein the throughshaft is arranged to extend through an internal annulus of the sun gear.14. The gas turbine of claim 1, wherein the ring gear is rigidlyconnected to the static front cone structure.
 15. The gas turbine ofclaim 1, wherein connection between the input shaft and the planet gearscomprises a friction locking spline connection.
 16. The gas turbine ofclaim 1, wherein the input shaft comprises flexibility means.
 17. Thegas turbine engine of claim 16, wherein the input shaft has a greaterflexibility than the fan shaft in at least bending and shear.
 18. Thegas turbine of claim 1, wherein fan shaft torsional stiff.
 19. The gasturbine of claim 1, wherein the gearbox device comprises an epicyclicgearbox with the ring gear being fixed relative to the other parts ofthe gearbox device and the output shaft device being connected to theplanet carrier.
 20. The gas turbine of at claim 1, wherein the gearboxdevice comprises a planetary gearbox in star arrangement with the planetcarrier fixed relative to the other parts of the gearbox device and theoutput shaft device being connected to the ring gear.